This invention relates generally to a gas turbine engine cooling and more particularly to cooling of airfoils such as turbine blades and nozzles.
High performance gas turbine typically rely on increasing turbine inlet temperatures to increase both fuel economy and overall power ratings. These higher temperatures, if not compensated for, oxidize engine components and decrease component life. Component life has been increased by a number of techniques.
Many solutions to improved components involve changing materials used in fabricating the components. U.S. Pat. No. 653,579 issued to Glezer et al on Aug. 5, 1997 shows a turbine blade made of a ceramic material. Other systems instead use a coating to protect a metal turbine blade as shown in U.S. Pat. No. 6,039,537 issued to Scheurlen on Mar. 21, 2000.
Even improved materials typically require further cooling. Most components include a series of internal cooling passages. Conventionally, a portion of the compressed air is bled from an engine compressor section to cool these components. To maintain the overall efficiency of the gas turbine, only a limited mass of air from the compressor section may be used for cooling. U.S. Pat. No. 5,857,837 issued to Zelesky et al on Jan. 12, 1999 shows an air foil having impingement jets to increase heat transfer. Impingement cooling creates high local heat transfer coefficients so long as spent cooling air may be effectively removed to prevent building a boundary layer of high temperature spent cooling air. Typically removal of spent cooling air is through a series of discharge holes located along the leading edge of the turbine blade. These systems require relatively high masses of cooling air. Further, plugging of the leading edge discharge holes may lead to a reduction of cooling and ultimately failure of the turbine blade.
Due to the limited mass of cooling air available and need to reduce pressure loss, component design requires optimal use of available cooling air. Typically, hot spots occur near a leading edge of a component. U.S. Pat. No. 5,603,606 issued to Glezer et al on Feb. 18, 1997 shows a cooling system that induces vortex flows in the cooling fluid near the leading edge of the component to increase heat transfer away from the component into the cooling fluid. The cooling flow in this system is limited by the size of the downstream openings in the turbine blade or component.
The present invention is directed to overcome one or more of the problems as set forth above.
In one aspect of the current invention an air foil has a leading edge and trailing edge. A first gallery is disposed internally in the air foil near the leading edge. A second radial gallery is disposed between a peripheral wall of the air foil and the first gallery. The second gallery is in fluid communication with the first gallery. A film cooling gallery is disposed internally of the peripheral wall proximate the leading edge. The film cooling gallery is fluidly connected with the second gallery and has a plurality of openings extending through the peripheral wall.
In another aspect of the present invention a method of cooling an air foil requires supplying a first portion of cooling fluid through a plurality of holes into a gallery adjacent an inner surface of a peripheral wall proximate a leading edge of a air foil. A film portion of the first portion of cooling fluid is transferred to a film cooling gallery. The film cooling gallery is connected to an outer surface of the peripheral wall near the leading edge (150).